1. Field of the Invention
The present invention relates to devices that provide vibration attenuation and isolation at structural joints and attachments.
2. Description of Related Art
Mechanical vibration is a term that describes oscillatory motion resulting from fluctuating forces acting on a dynamic system, that is, a system of mass and elastic elements. In certain situations, these motions can become excessive, causing reduced life, improper function, and possible failure of the system. This is especially important in regard to aircraft, or more specifically, rotorcraft structures, where failure of the structure may result in loss of life or aircraft. Excessive vibration within these structures may also lead to discomfort or sickness of passengers or crew, impairing safe operation of the aircraft. Effective control of vibrations is very important in this and other applications.
Of primary interest is the resonant condition, where masses and elastic members vibrate at or near their natural frequency. Referring to FIGS. 1A and 1B in the drawings, a simple dynamic system 11 where resonance can occur is illustrated schematically and with a plot, respectively. As is shown, a mass 13 is supported by a spring 15 and a damper 19 attached to a moving support 17. The motion of support 17 is oscillatory in the vertical direction. For this system, the natural frequency is simply the frequency at which mass 13 would oscillate if displaced and let go, with support 17 not moving. When support 17 is moving, the resulting motion of mass 13 with respect to support 17 depends upon the amplitude and frequency of the motion of support 17. If the driving frequency of support motion f becomes equal to the system natural frequency fn, resonance occurs, which results in very large motions of mass 13 for systems that are lightly damped. This is typical of many components and structures, and is illustrated on a plot 21 in FIG. 1B, where the maximum motion X of mass 13 with respect to the motion Y of support 17 occurs when f/fn=1.
For dynamic systems in general, a resonant condition is undesirable and potentially destructive, and should be avoided. This can be accomplished by controlling the driving frequency f and/or the natural frequency fn, or by incorporating sufficient damping. For many systems, such as helicopters, the driving frequency f remains almost constant, i.e., rotor at constant RPM, and sufficient damping is hard to implement without additional weight. As a result, avoiding resonance requires controlling the system natural frequency fn, so that the natural frequency fn is never equal to the driving frequency f. This can be done by either changing the mass or the stiffness properties of the system. Because the mass is usually fixed, the only remaining adjustment is the stiffness of the system.
The simple model of FIGS. 1A and 1B can be extended to more complex systems having multiple mass and stiffness elements, such as helicopter or tilt-rotor airframes. For these systems, multiple driving frequencies from the vibrating rotor combined with the distributed mass and stiffness throughout the airframe create a complex problem in vibration control. Historically, this problem has been overcome by isolating the structure from vibrating components, i.e., rotors, pylons, etc., and/or by building the structure very stiff, so that the system natural frequencies remain higher than any driving frequencies in the system. These frequency isolation methods are simple, but cannot be incorporated without adding significant weight to the airframe structure.
Current efforts to extend the state of the art involve the development of dynamically tailored airframe structures that are “adaptive,” or able to change their dynamic characteristics as desired. By changing the stiffness properties of the structures, the structures are able to “de-tune” themselves from adverse resonant conditions, allowing less stiff and potentially lighter structures.
In absence of an effective and practical means to change the stiffness of elastic members in dynamic systems, vibration is often controlled by isolating vibrating components. In a broad context, isolation simply means allowing the vibrating components to move independently, as much as possible, in such a way as to minimize transmitted forces to the remaining system. This type of vibration control is commonly done by supporting or connecting the vibrating components with flexible elements. As applied to a helicopter, a common method of achieving this is by supporting the fuselage from the vibrating rotor and pylon using elastomeric supports acting as springs.
A model of a simple helicopter dynamic system 20 is illustrated in FIGS. 2A and 2B in the drawings. As is shown, a rotor/pylon 23 is supported by a spring 25 and a damper 29 attached to a fuselage 27. The frequency response of system 20 is shown in a plot 31 of FIG. 2B. As shown in the far right of plot 31, the relative motion X of fuselage 27 with respect to the motion Y of rotor/pylon 23 becomes small when the natural frequency fn is much smaller than the frequency of motion f of rotor/pylon 23. In this system, the softer the spring 25, the lower the natural frequency fn, and corresponding motion X of fuselage 27. With regard to vibration isolation, any support in such an application should be as soft as possible. However, if spring 25 is too soft, excessive deflections can occur as rotor loads change.
U.S. Pat. No. 4,362,281 issued to Cresap et al. is based upon this principle, and embodies a soft spring support for isolation of the vibrating rotor/pylon during steady flight conditions. To prevent excessive deflections during changing flight conditions and variations in rotor thrust, mechanical stops are incorporated that “bottom out” and limit motion during these transient conditions. Thus, in the Cresap et al. system, the system stiffness changes from relatively soft to effectively very stiff at the limits of pylon motion.
In some dynamic helicopter systems, dynamic components themselves are used as supports between the helicopter rotor/pylon and the fuselage. The dynamic antiresonant vibration isolator (DAVI) is an example of such an approach. A simple model of DAVI system 41 is illustrated in FIG. 3. In DAVI system 41, a fuselage mass 43 is attached to a rotor/pylon 45 using a spring element 47 in parallel with a weight 49 on a lever 51. The mechanical advantage of weight 49 and lever 51 can be tailored so that when rotor/pylon 45 is oscillating at a particular frequency, the inertial and spring forces acting on fuselage 43 through a lever pivot 53 and spring element 47 are equal and opposite, so that, theoretically, no net forces are acting on the fuselage.
The devices disclosed in U.S. Pat. No. 6,247,684 issued to Manfredotti and U.S. Pat. No. 4,365,771 issued to Halwes are based upon the DAVI principle. Manfredotti discloses a dynamic component intended for use as a support between a helicopter rotor/pylon and fuselage, and Halwes discloses a liquid inertia vibration isolator. In these devices, the net forces acting on the structure are minimized, thereby limiting vibration. These devices, however, are only effective within a narrow frequency band of operation, and may not provide adequate isolation as rotor rpm, flight, or operating conditions change.
The rotary beam variable stiffness wing spar described in U.S. Pat. No. 6,000,660 issued to Griffin et al. discloses a variable stiffness element for use in dynamically tailored airframe structures. In the Griffin et al. device, the wing spar is a non-rectangular beam, having different bending stiffness depending upon the orientation of the beam and loading. When rotated within the wing, the stiffness and dynamic properties of the wing can be varied. The Griffin et al. device is very large and heavy, and difficult or impractical to implement in but a few locations.
Although the foregoing designs represent considerable advancements in the area of vibration isolation and control, many shortcomings remain.